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GOE 6K AIRFOIL (goe06k-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 6K AIRFOIL (goe06k-il)
Reynolds number: 100,000
Max Cl/Cd: 45.65 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe06k-il-100000.txt
Download as CSV file: xf-goe06k-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 6K AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.4860   0.11086   0.10608  -0.0337   1.0000   0.0588
  -8.500  -0.5003   0.10919   0.10449  -0.0333   1.0000   0.0589
  -8.250  -0.5136   0.10721   0.10258  -0.0333   1.0000   0.0589
  -8.000  -0.4782   0.09884   0.09415  -0.0276   1.0000   0.0632
  -7.750  -0.4857   0.09658   0.09195  -0.0260   1.0000   0.0644
  -7.500  -0.4972   0.09454   0.08998  -0.0242   1.0000   0.0664
  -7.250  -0.5078   0.09221   0.08771  -0.0232   1.0000   0.0671
  -7.000  -0.5168   0.08948   0.08503  -0.0241   1.0000   0.0699
  -6.750  -0.5218   0.08649   0.08205  -0.0254   1.0000   0.0709
  -6.500  -0.5263   0.08387   0.07929  -0.0301   1.0000   0.0723
  -6.250  -0.5269   0.07890   0.07441  -0.0281   1.0000   0.0742
  -6.000  -0.5237   0.07587   0.07144  -0.0248   1.0000   0.0768
  -5.750  -0.5199   0.07277   0.06830  -0.0244   1.0000   0.0801
  -5.000  -0.4992   0.06211   0.05737  -0.0255   1.0000   0.0923
  -4.750  -0.4853   0.05880   0.05362  -0.0277   1.0000   0.1014
  -4.500  -0.4768   0.05543   0.05035  -0.0256   1.0000   0.1053
  -4.250  -0.4629   0.05227   0.04693  -0.0259   1.0000   0.1169
  -4.000  -0.4491   0.04957   0.04403  -0.0254   1.0000   0.1303
  -3.750  -0.4354   0.04695   0.04130  -0.0245   1.0000   0.1446
  -3.500  -0.4220   0.04450   0.03884  -0.0231   1.0000   0.1615
  -3.000  -0.3994   0.04073   0.03510  -0.0193   1.0000   0.2348
  -2.250  -0.2974   0.02963   0.02114  -0.0176   1.0000   0.0836
  -2.000  -0.2732   0.02739   0.01825  -0.0157   1.0000   0.0749
  -1.750  -0.2520   0.02604   0.01674  -0.0145   1.0000   0.0789
  -1.500  -0.2291   0.02479   0.01519  -0.0131   1.0000   0.0807
  -1.250  -0.2059   0.02373   0.01388  -0.0118   1.0000   0.0820
  -1.000  -0.1831   0.02293   0.01286  -0.0105   1.0000   0.0872
  -0.750  -0.1615   0.02210   0.01209  -0.0095   1.0000   0.0975
  -0.500  -0.1404   0.02133   0.01135  -0.0081   1.0000   0.1064
  -0.250  -0.0427   0.01835   0.01129  -0.0223   1.0000   1.0000
   0.000  -0.0219   0.01866   0.01118  -0.0211   1.0000   1.0000
   0.250  -0.0016   0.01899   0.01126  -0.0202   1.0000   1.0000
   0.500   0.0187   0.01934   0.01140  -0.0193   1.0000   1.0000
   0.750   0.0390   0.01971   0.01161  -0.0184   1.0000   1.0000
   1.000   0.0591   0.02009   0.01187  -0.0176   1.0000   1.0000
   1.250   0.0792   0.02050   0.01218  -0.0169   1.0000   1.0000
   1.500   0.0992   0.02093   0.01253  -0.0161   1.0000   1.0000
   1.750   0.1190   0.02139   0.01291  -0.0154   1.0000   1.0000
   2.000   0.1388   0.02187   0.01334  -0.0147   1.0000   1.0000
   2.250   0.1841   0.02285   0.01430  -0.0194   0.9904   1.0000
   2.500   0.2261   0.02379   0.01525  -0.0232   0.9818   1.0000
   2.750   0.2661   0.02459   0.01608  -0.0266   0.9730   1.0000
   3.000   0.3005   0.02512   0.01666  -0.0288   0.9629   1.0000
   3.250   0.3360   0.02569   0.01732  -0.0312   0.9528   1.0000
   3.500   0.3735   0.02628   0.01804  -0.0338   0.9428   1.0000
   3.750   0.4155   0.02687   0.01877  -0.0371   0.9328   1.0000
   4.000   0.4569   0.02726   0.01933  -0.0402   0.9209   1.0000
   4.250   0.5065   0.02692   0.01922  -0.0438   0.8984   1.0000
   4.500   0.6397   0.02261   0.01557  -0.0576   0.8515   1.0000
   4.750   0.7291   0.01597   0.00950  -0.0591   0.7405   1.0000
   5.000   0.7536   0.02012   0.00967  -0.0565   0.1077   1.0000
   5.250   0.7661   0.02149   0.01094  -0.0535   0.0870   1.0000
   5.500   0.7886   0.02333   0.01273  -0.0522   0.0767   1.0000
   5.750   0.8258   0.02572   0.01505  -0.0539   0.0664   1.0000
   6.000   0.8644   0.02805   0.01755  -0.0556   0.0604   1.0000
   6.250   0.9049   0.03131   0.02103  -0.0574   0.0585   1.0000
   6.500   0.9371   0.03600   0.02602  -0.0581   0.0568   1.0000
   6.750   0.9519   0.03725   0.02782  -0.0547   0.0550   1.0000
   7.000   0.9762   0.04196   0.03280  -0.0538   0.0572   1.0000
   7.250   1.0010   0.04758   0.03990  -0.0476   0.0986   1.0000
   9.250   0.9732   0.07981   0.07468  -0.0171   0.1350   1.0000
   9.500   0.9083   0.07876   0.07390  -0.0101   0.1365   1.0000
   9.750   0.8732   0.08059   0.07587  -0.0049   0.1370   1.0000
  10.000   0.8393   0.08400   0.07936  -0.0020   0.1371   1.0000
  10.250   0.8065   0.08850   0.08392  -0.0014   0.1372   1.0000
  10.500   0.7714   0.09414   0.08959  -0.0031   0.1374   1.0000
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